Abstract: The expense of fabricating an annular combustor (10) for a gas turbine is minimized by providing a combustor housing (12) including a generally axially extending sleeve (14) and an annular liner (18) disposed within the housing (12) and about the sleeve (14). The annular liner (18) has concentric inner and outer axially extending walls (20, 22) spaced from the sleeve (14 ) and the housing (12), respectively, and also has a radially extending wall (24) spaced from the housing (12 ) and interconnecting the inner and outer walls (20, 22) at one end to define a combustion chamber (26). The liner (18) is spaced from the housing (12 ) and the sleeve (14 ) to define a compressed air flow path (28) extending from a radially outer compressed air inlet (30) to a radially inner compressed air outlet (32) in communication with the combustion chamber (26) axially remote from the radially extending wall (24).

Abstract: A cylindrical combustion chamber housing of a gas turbine, in which the compressor air is fed into the lower, conical part of the combustion chamber housing, the perforated cone (5), through a lateral, arc-shaped inlet elbow (2). The inlet elbow (2) is directly joined by the intake distribution element (1), in which the compressor air is led around the perforated cone (5) on both sides. The tangential flow is converted around a cone into an axial flow through the holes (7) in the perforated cone (5). The conversion of the direction of flow of the compressor air is supported by radially arranged ribs (3). As a result, optimal cooling of the entire injector tube (6) is achieved, while the pressure drop in the air feed area (1, 2 and 5) is minimized, and the efficiency of the gas turbine is increased at the same time.

Abstract: A premixed combustor for a gas turbine, comprises, a combustion chamber for burning a fuel with an air therein, at least two premixed fuel/air mixture outlets which are juxtaposed with each other, and through each of which a premixed fuel/air mixture flows out into the combustion chamber, and a flow deflection member arranged between the premixed fuel/air mixture outlets, urging the permixed fuel/air mixture from one of the premixed fuel/air mixture outlets to go away from another one of the premixed fuel/air mixture outlets, and terminating in the combustion chamber so that the premixed fuel/air mixture from the one of the premixed fuel/air mixture outlets is permitted to move toward the premixed fuel/air mixture flowing out from the another one of the premixed fuel/air mixture outlets after going away from the another one of the premixed fuel/air mixture outlets.

Abstract: A hot gas generator includes a dual fuel injector (42) having spaced fuel discharge nozzles (48, 60) on the longitudinal axis (28) of a vessel (18) having a narrow inlet (22), an opposite narrow outlet (14) and an intermediate, enlarged chamber (24) which serves as a combustion chamber. The nozzle (60) injects a spray cone (64) of fuel in a cone angle that is sufficiently wide that a large angle exists between the fuel spray and the path (54) of high velocity combustion gases moving toward the outlet (14) to achieve excellent vaporization of fuel at fuel rich conditions without causing carbon buildup and/or the generation of black smoke in the exhaust.

Abstract: Improved fuel atomization for turbine engines utilizing high viscosity fuels is accomplished in an engine having an annular combustor (26) by providing an air injector (52) which produces an air stream (50) generally tangentially directed into the combustor (26) to intersect a fuel spray or film (46) also generally tangentially directed into the combustor (26) by a fuel injector (48) to atomize fuel for ignition by an igniter (60) in an annular outer wall (34) of the combustor (26).

Abstract: A reverse flow annular combustor is provided having downstream wall comprised of a shroud covering the outer surface of the outer transition liner. The shroud abuts the transition liner at a plurality of points to define a plurality of passages between the shroud and the liner. The passages receiving cooling through the shroud at its radial inner end and expel the cooling air through the liner at its outer radial portion and into the hot gas in the combustor. Thus, the air used to cool the liner is also used for dilution mixing with the combustion gas.

Abstract: To control cooling by oxidant surrounding a combustion chamber (20, 20'), a stored energy combustor (10, 10') includes an oxidant flow path (28) which extends about the combustion chamber (20, 20'). The stored energy combustor (10, 10') comprises a vessel (12, 12') having a wall (14, 14') defining spaced inlet and outlet ends (16 and 18) interconnected by the combustion chamber (20, 20'), an oxidant inlet port (22) upstream of the combustion chamber (20, 20') for directing oxidant thereinto, and a fuel injector (24) upstream of the combustion chamber (20, 20') for directing fuel thereinto. A housing (26) substantially surrounds the vessel (12, 12') in spaced relation to the wall (14, 14') from the inlet end (16) to the outlet end (18) whereby the housing (26) and wall (14, 14') define the oxidant flow path (28) therebetween which is in communication with the oxidant inlet port (22) upstream of the combustion chamber (20, 20').

Abstract: Difficulties in cooling the shrouds (32, 34) and the nozzle (38) of an air breathing gas turbine are minimized or eliminated through the use of a generally toroidal, annular combustor (44) having an outlet (48) and combustion air inlets (90A, 90B) and otherwise being imperforate or free of any means for the admission of cooling air to the interior of the combustor (44).

Abstract: In order to establish a larger dome height to thereby increase combustion volume within the same turbine engine envelope, the inner wall region of an annular combustor is formed to define a thrust nozzle which in turn provides an enhanced combustion zone. The turbine engine will include a rotary compressor and a turbine wheel coupled to the rotary compressor for driven movement thereof along with an annular nozzle proximate the turbine wheel for directing gases of combustion thereat. The annular combustor will be disposed about the turbine wheel and defined by a radially extending wall region connecting the inner wall region to an outer wall region such that the annular combustor has a combustion zone upstream of the annular nozzle and a dilution air zone intermediate the annular nozzle and the combustion zone.

Abstract: In order to facilitate operation in a wide variety of operating conditions, and to enhance altitude starting capability while eliminating start injectors, a radial turbine engine (10) includes dual fuel injection zones (24 and 26). The radial turbine engine (10) also includes a turbine wheel (12) coupled to a rotary compressor (14) for axially driven movement, an annular nozzle (16) for directing gases of combustion radially at the turbine wheel (12), and an annular combustor (15). The annular combustor (15) defines an annular combustion space disposed about the turbine wheel (12) and in fluid communication with both the compressor (14) and the nozzle (16), and it receives fuel from a source and air from the compressor (14) which are combusted in the combustion space to generate the gases of combustion.

Abstract: In order to avoid plugging the air passageways (70) of fuel injectors (42) with carbon particles in a gas turbine engine (10), the gas turbine engine (10) includes an annular combustor (18) having radial dilution air injection. The gas turbine engine (10) also includes a rotor (12) having turbine blades (14) and a nozzle (16) adjacent the turbine blades (14) which is adapted to direct hot gases at the turbine blades (14) to cause rotation of the rotor (12). The annular combustor (18) is disposed about the rotor (12) and has an outlet (20) to the nozzle (16), spaced inner and outer walls (22 and 24), and a generally radially extending wall (26) connecting the inner and outer walls (22 and 24). The gas turbine engine (10) further includes a housing (28) substantially surrounding the annular combustor (18) in spaced relation to the inner, outer, and radially extending walls (22, 24 and 26) to define a dilution air flow path (30).

Abstract: A combustion structure for a turbine engine includes an annular combustion cavity into which flows a part of each of a plurality of radially outwardly flowing primary air jets to define a rotating toroid of primary combustion air. Plural annular louvers defined in the walls of the combustion cavity provide plural film-like wall cooling air flows moving in agreement with the rotating toroid of primary combustion air. An annular array of circumferentially spaced apart and disposed louvers in a radially outer wall of the combustor structure provide film-like air flows to energize the rotating toroid of combustion air to swirl about the combustion cavity.

Abstract: Difficulties in accommodating thermal growth in a combustor for providing hot gases of combustor to drive, for example, a turbine wheel (10) are avoided in a combustor construction, including a housing (36) defining a combustion chamber having a relatively narrow inlet (38), an enlarged combustion area (42) and a relatively narrow outlet (40) opposite of the inlet (38). A fuel injector (46) is disposed in the inlet (38) for injecting fuel (50) at least into the combustion area (42) and a plenum case (56) surrounds the housing (36) in generally spaced relation. The plenum case (56) is adapted to be connected to a source (62) of oxidant under pressure and is in fluid communication with the inlet (38) to provide oxidant thereto to support combustion of fuel in the combustion area (42). A conduit (44) extends from the outlet (40) to be connected to a turbine nozzle (16) and a diaphragm (122) is interposed and sealingly engages both the housing (36) and the plenum case (56) near the outlet (40).

Abstract: Problems with the cooling of a turbine nozzle (60) and a turbine wheel (20) of a gas turbine engine when the engine is uprated may be avoided or minimized by employing a combustor (36) that is free of any means for providing cooling air films on the interior walls thereof and which allows the entry of only stoichiometric quantities of air into the interior of the combustor (36) at predetermined locations along the entire axial length thereof. Consequently, the entirety of the combustor (36) is available for combustion. Cooling of the nozzle (60) and components of the engine downstream thereof is handled by the provision of an annular opening (96) immediately upstream of the leading edges (98) of the vanes (58) constituting the nozzle (60) along with apertures (102) in a front turbine shroud (30) and aligned with leading edges (98). Air from the engine compressor (18) is flowed through the apertures (102) and the opening (96).

Abstract: A gas turbine engine is disclosed having a mixed-flow compressor and a connecting mixed-flow diffuser. The mixed-flow diffuser is disposed inside the annularly arranged combustion chambers and supplying them, from the direction of the radial interior side, in the reverse flow, with compressor air. Gas turbine engines of this type are more compact and show a higher efficiency than turbine engines having radial flow compressors.

Abstract: In order to significantly reduce the carbon produced in a combustor (10), and thus reduce or eliminate carbon buildup on combustor walls (14), a stored energy combustor (10) has a fuel discharge port (30) downstream of the combustion chamber (20). The downstream fuel discharge port (30) directs fuel into a secondary stored energy zone (24) downstream of a primary stored energy zone (22) comprised of an inlet end (16) and the combustion chamber (20) which receives an oxidant and a fuel through an upstream fuel discharge port (28) where the oxidant/fuel mixture is ignited to produce hot gases of combustion. The downstream fuel discharge port (30) is arranged to direct fuel in a manner forming a fuel film on the interior wall (14) in the secondary stored energy zone (24) such that the hot gases of combustion interact with the fuel film to cause fuel evaporation by establishing a stratified fuel annulus (54, 56).

Abstract: The expense of a cover (80) for non-uniform structure (38) on or adjacent a surface (90) of a rear turbine shroud 34 facing a cooling passage defined by the rear engine shroud (34) and a wall (62) of a combustor (50) in a gas turbine engine can be avoided by providing a connecting wall (100) that is smoothly curved between the radially inner wall (58) and a conical wall (62) of the combustor (50). The resulting cooling air stream will attach itself to the walls (100) and (62) by reason of the Coanda effect, rendering the presence of non-uniform structure (38) on the surface (90) unimportant.

Abstract: In order to significantly reduce the carbon produced in a combustor, and thus reduce or eliminate carbon buildup on a combustor wall (14), a stored energy combustor (10) has a fuel injector (28) in a tubular extension (16) of a vessel (12) leading to a combustion chamber (20). The fuel injector (28) has a discharge end (60) upstream of a plane defining a point of entry (32) into the combustion chamber (20) for directing an annulus of fuel (26) toward the combustion chamber (20), and an oxidant inlet port (24) is also provided upstream of the combustion chamber (20) for directing oxidant into the combustion chamber (20) through the tubular extension (16) in surrounding relation to the fuel injector (28) which is concentric with the longitudinal axis (22) of the vessel (12).

Abstract: Inexpensive cooling of a gas turbine nozzle and shroud assembly 62, 66, 76 is obtained by locating passages 84, 88, 102 in the vanes 76 forming part of the nozzle structure 62, 66, 76 so as to open generally centrally of the vanes 76 in the area between the shrouds 62 and 66 and in the leading edge of the vanes 76.

Abstract: An annular combustor having radial dilution air injection in a gas turbine is disclosed. The gas turbine includes a rotor having turbine blades and a nozzle adjacent the turbine blades which is adapted to direct hot gases at the turbine blades to cause rotation of the rotor. The annular combustor is disposed about the rotor and has an outlet to the nozzle, spaced inner and outer walls, and a generally radially extending wall connecting the inner and outer walls. The gas turbine includes a housing substantially surrounding the annular combustor in spaced relation to the inner, outer and radially extending walls to define a dilution air flow path. The dilution air flow path includes a compressed air inlet in communication with a compressor supplying dilution air at one end thereof and a compressed air outlet in communication with the annular combustor adjacent the outlet at the other end thereof.

Abstract: A gas turbine engine 10 having a rotor 12 with turbine blades 14 and a nozzle 16 adjacent the turbine blades 14 for directing hot gases of combustion at the turbine blades 14. The engine 10 also includes an annular combustor 18 about the rotor 12 and an annular combustor housing 30 substantially surrounding the combustor 18 in generally concentric spaced relation. A fuel injection system is operatively associated with the combustor 18 for injecting fuel into a combustion space 28 to be combusted with air from a compressor 34. The fuel injection system may include a plurality of circumferentially spaced fuel injectors 68 disposed in an outer wall 20 of the combustor 18 together with a generally oval shaped manifold 70 in fluid communication with a primary fuel source and the fuel injectors 68. A turbine shroud 36 is provided to extend radially outward from the rotor to the outer wall of the combustor on the side of the nozzle 16 opposite the combustion space 28.

Abstract: Unequal flow rates of dilution air about an annular combustor and the resulting hot spots and thermal stresses are avoided in a structure wherein a vane assembly (66) made up of ring (76) and integral vanes (76) is mounted to the rear turbine shroud (42) for limited movement with respect thereto sufficient to accommodate unequal thermal growth of the rear turbine shroud (42) and the vane assembly (66).

Abstract: The cooling and mounting of vanes 50 in a turbine nozzle 24 between shrouds 28, 29 is facilitated by impaling the vanes 50 with threaded fasteners 54, 70 and conducting relatively cool air from the compressor 15 through the grooves 72 between adjacent threads 70 of the threaded fastener 54.

Abstract: Improved performance in a hot gas generator is achieved by disposing a dual fuel injector 42 having spaced fuel discharge ports 48, 60 on the longitudinal axis 28 of a vessel 18 having a narrow inlet 22, an opposite narrow outlet 14 and an intermediate, enlarged chamber 24 which serves as a combustion chamber.

Abstract: The large mass of a containment ring for a gas turbine and the cost of the material fabricating the same may be minimized in a gas turbine including a rotor 12 having a radial inflow turbine wheel with blades 14 and a nozzle 50 adapted to direct hot gases at the blades 14 to cause rotation of the rotor 12. An annular combustor 18 for supplying gases to the nozzle 50 is surrounded on its inner and outer sides by a housing including a rear shroud 46 and in turn defines a dilution air outlet 34 just upstream of the nozzle 50. A containment ring 70, 80, 90 is disposed in the dilution air path just upstream of the nozzle 34 so as to be cooled by air flowing therein. As a consequence, the containment ring 70, 80, 90 may be located at a radially inward position to minimize its mass and is adequately cooled so as to allow the use of nonexotic materials in fabricating the same.

Abstract: Improved performance in a hot gas generator 10 is achieved by providing a pair of hemispherical liners 22 and 24 loosely positioned within an interior wall 14 of a vessel 12 so as to be disposed above a combustion chamber 20 therein. The hemispherical liners 22 and 24 normally have a groove 26 at an interface therebetween and are formed of a material adapted to thermally expand under heat. Specifically, the hemispherical liners 22 and 24 expand to close the groove 26 at the interface in a manner producing relatively little stress thereon.

Abstract: Undesirably high and damaging temperature gradients resulting from hot spots in combustors 26 for housing a hot gas generating oxidation reaction are avoided in a construction including a combustor housing haivng a wall 32, 34, 39 with an interior surface defining a combustion space 38, 40 and an exterior surface thereof and provided with an outlet 36. A plenum 44, 80, 82 surrounds the combustor 26 and a fuel injector 50, 52 is provided for introducing a fuel to be oxidized into the combustion space 38, 40. Oxidant inlets 54 to the combustion space 38, 40 are provided and various structures including the plenum 44, 80, 82 flow cooling gas in a path about the exterior surface of the walls, 32, 34, 39 to cool the combustor housing. Trip strips 114, 118 are located on the exterior surface of the walls 32, 34, 39 and extend into the cooling gas flow path and are located to minimize the temperature gradient between points along those walls 32, 34 and 39.

Abstract: Reliable starting of air breathing turbines even at high altitude is assured in a turbine engine of the type including an annular combustor 26 by locating one or more start injectors 96 away from the bottom of the combustor 26 and making the bottom of the combustor 26 free of start injector 96.

Abstract: The clearance between a seal assembly 48 and a turbine wheel 16 in an engine of the type having a radial outflow compressor 24 and a radial inflow turbine 16 is minimized by forming the seal assembly 48 in part out of a plurality of segments 74 disposed in a circular array and which are relatively movable but sealed to each other. The thermal stress in the segments 74 will be less than in a single ring resulting in an extended life before the onset of cracking and eventual mechanical failure.

Abstract: In a staged low NOx lean premix hot wall gas turbine combustor, the interior wall of the combustor is maintained at or near the flame temperature to provide stable operation over a wide range of heat release values. One or more stages of the staged combustor includes a Stirred Reactor region followed by a Plug Flow Reactor region. In one embodiment, a lean premixture of fuel and air is inducted tangentially into an annular intake manifold for inducting a cyclonic flow of the premixture into the combustion chamber. In another embodiment, the Stirred Reactor region is developed by colliding, preferably head-on, a plurality of jets of lean premixture within the combustion chamber.

Abstract: The combustion dynamics and efficiency of a gas turbine having an annular combustor 26 provided with fuel injection nozzles 50 that inject fuel generally tangentially is improved by providing the walls 32, 34, 39 of the combustor 26 with cooling air film injectors 70, 86; 72, 88; 74, 90 at substantially equally angularly spaced locations about each such wall and which are oriented to generally tangentially inject a film-like air stream on the associated wall 32, 34, 39.

Abstract: Warping of the rear turbine shroud 54 and the resulting cracking of turbine nozzle vanes 24 may be eliminated in a radial inflow turbine having an annular combustor 30 with a radially outer wall 32 opening on a compressed air plenum 48 by locating a plurality of tubes 70 on the radially outer wall 32 in fluid communication with the plenum 48 and extending the tubes 70 inwardly so that the radially inner ends 76 are adjacent the radially outer edge 56 of the turbine shroud 54. The tubes 70 thus direct sweeping streams of cooling air along the rear turbine shroud 54 to cool the same and prevent warpage.

Abstract: In a gas turbine the transition duct is subject to high temperature since it conducts the combustion products to the power turbine. It is normally cooled by air from the compressor. This invention improves the uniformity of cooling by forcing the cooling air to pass over all surfaces of the transition duct. Air flow passages are formed between the ducts and saddle shaped members which direct the flow of air from the compressor over the outer surfaces of the ducts.

Abstract: The combustion chamber according to the present invention utilizes a double-walled structure only in the converging portion of the combustion chamber. A hot wall is retained adjacent to, but spaced from a corresponding cold wall so as to provide a cooling space between them. An upstream portion of the hot wall is retained in a notch defined by the cold wall, while the downstream portions each have generally radially extending flanges. Studs pass through openings in the adjacent flanges to prevent any relative circumferential movement, while a clamp serves to attach the flanges to the inlet of a turbine. The clamp slidably retains the flanges therein so as to permit relative radial movement to accommodate any thermal expansion and contraction of the combustion chamber, solely by the appended claims.

Abstract: A shield is disposed in spaced relationship to the exterior surface of the combustor liner of a gas turbine engine and extends over a plurality of very small holes formed in the combustor liner so as to block flow of air to these holes. Further, the shield provides a space between the shield and the combustor liner for receiving reverse flow of air. A portion of the air flowing along the exterior surface of the combustor liner is diverted for reverse flow in the path provided between the shield and the combustor liner. A lesser plurality of substantially larger holes are provided in the combustor liner at the point of reversal of the air so that the dirt particles in the air, because of the centrifugal force acting thereon, tend to flow through these larger holes. A second group of larger holes are provided in the combustor liner approximately at the forward end of the space between the shield and the combustor liner.

Abstract: The cost of fuel injection nozzles 50 and their tendency to clog in a gas turbine having an annular combustor 26 can be reduced by alternating the fuel injection nozzles 50 with bender jets 56 configured to introduce a combustion supporting gas into an annular combustion zone 40 at locations between the fuel injectors 50 to achieve uniform turbine inlet temperature distribution while requiring fewer of the nozzles 50 and allowing those nozzles 50 that are utilized to have larger fluid flow paths that are less prone to clogging.

Abstract: A gas turbine combustion chamber with an inlet and outlet has apertures in the combustion chamber wall and air scoops provided in the apertures, the air scoops having an outer tubular member and coaxially spaced inner tubular member. The spaced tubular members have inner cylindrical portions and outwardly extending flanges, with spacers disposed between and secured to the flanges with cooling air passing therebetween and through the annular air flow passage into the combustion chamber. Preferably, a radially outwardly extending arcuate section on the inner tubular member has a radius at least about one-third of the inner diameter of the inner cylindrical portion thereof.

Abstract: An air admission insert particularly adapted for air admission holes in a combustor liner in a gas turbine combustion system comprises a pair of sleeve members, one of which is inserted in the other in axial but eccentric relationship so that the sides of the sleeves contact each other with the intervening space between the sleeves being crescent shaped. The insert is positioned concentrically in a combustor liner air admission hole. Combustion air flows through the insert and into the combustor liner and axially through the crescent space to cool the insert as well as the liner hold perimeter.

Abstract: A gas turbine cooling apparatus is disclosed in which a flow sleeve is provided to surround a combustor liner and a tail pipe substantially over their full length. A group of small holes for impinge-cooling an outer wall of the tail pipe are formed in a region of the flow sleeve close to a turbine. Further, opening portion for introducing cooling air are provided closer to the combustor liner than the small holes. Thus, the outer wall of the tail pipe and the wall of the combustor liner are cooled by the cooling air flowing between the tail pipe, the combustor liner and the flow sleeve.

Abstract: Aircraft engine cooling apparatus having a plurality of cooling gas inlet nozzles, producing first streams of cooling gas moving in a forward direction, a plurality of first direction reversal members for splitting the first streams into second and third streams straddling the nozzles, and substantially reversed in direction relative to the direction of the first streams, a plurality of second direction reversal members, positioned between pairs of adjacent nozzles, for again substantially reversing the directions of the second and third streams, while combining the second and third streams to form a fourth stream of laminar flow cooling film, and wherein adjacent pairs of the first direction reversal members have wall portions configured to form gradually diverging channels through which the fourth stream cooling film moves.

Abstract: A combustion chamber liner having a layer of refractory material, a layer of metallic material, means to adaptably fix the layer of refractory material to the layer of metallic material and a means for internally cooling the liner. The means to adaptably fix the layer of refractory material to the layer of metallic material allows both layers to expand and contract separately without risk of damaging the liner because of different thermal expansion rates. The means for internally cooling the liner allows the liner to operate at higher temperatures and increases the useful life of the liner.

Abstract: A combustion chamber liner has three layers; a first ceramic layer, a second filamentary steel wool type layer and a third metallic layer. The interior of the liner is cooled by passages in the third layer and emergency or failure activated cooling occurs in a damaged area in the first layer. Sensor means can be placed in the liner or externally of the liner to detect changes in the coolant flow or pressure and thereby detect a failure in the liner.

Abstract: The expense of fabricating an annular combustor for use with a small diameter gas turbine engine is minimized in a construction including a bell-shaped housing 52, an axially extending sleeve 66 within the housing 52, an annular liner 74 within the housing 52 and about the sleeve 66 and including inner, outer and radial walls 80, 76, 78, respectively, spaced from the sleeve 66 and the housing 52 and all being formed of sheet metal. A circular fuel manifold 108 with angularly spaced fuel dispensing openings 122 is disposed between the housing 52 and the outer wall 76 and a plurality of open ended, elongated tubes 112 extend through the outer wall 76 and are in fluid communication with respective openings 122 in the fuel manifold 108 to introduce air and fuel generally tangentially into the liner 74. An end 68 of the sleeve 66 and an end 86 of the outer wall 76 define an outlet in fluid communication with a turbine nozzle structure 46 having turbine nozzle blades 136.

Abstract: A combustion chamber for a turbojet engine is disclosed which utilizes double walled construction. Hot walls are attached to adjacent cold walls solely by an orifice member which defines an orifice to supply primary of dilution air to the combustion chamber. The attachment allows relative radial movement of the hot walls with respect to the cold walls so as to accommodate the thermal expansion and contraction of the hot walls. The sleeves forming the walls also cooperate so as to provide an internal cooling film on the interior surface of the hot walls. Convection cooling may also be utilized between the walls to prevent structural damage due to high temperatures.

Abstract: The expense of fabricating an annular combustor for use with a small diameter gas turbine engine is minimized in a construction including a bell-shaped housing 52, an axially extending sleeve 66 within the housing 52, an annular liner 74 within the housing 52 and about the sleeve 66 and including inner, outer and radial walls 80, 76, 78, respectively, spaced from the sleeve 66 and the housing 52 and all being formed of sheet metal. A circular fuel manifold 108 with angularly spaced fuel dispensing openings 122 is disposed between the housing 52 and the outer wall 76 and a plurality of open ended, elongated tubes 112 extend through the outer wall 76 and are in fluid communication with respective openings 122 in the fuel manifold 108 to introduce air and fuel generally tangentially into the liner 74.

Abstract: A reverse flow combustion chamber includes an annular chamber enclosed between flame tube wall sections to which cooling air is so supplied from an outer annular channel acted upon with secondary air opposite the main flow direction in the flame tube in such a manner that the cooling air that it is blown out in the opposite flow direction film-like against an adjoining flame tube wall.

Abstract: A transition duct in an advanced heavy duty gas turbine engine is cooled by impingement jets formed by apertures in a sleeve spaced a distance from the surface to be cooled. The sleeve is configured so as to duct spent impingement air towards the combustor, where it can be subsequently used for mixing with, and combustion of, the fuel, or for cooling of the combustor. The distance between the impingement sleeve and the transition duct surface is varied to control the velocity of air crossflow from spent impingement air in order to minimize the pressure loss due to crossflow. The cross-sectional areas of the apertures are varied to project impingement jets over the various distances and crossflow velocities. Generally, larger aperture areas are used with larger distances.

Abstract: A gas turbine combustor has a plurality of burner units each having an inner tube defining a combustion chamber therein and an outer tube surrounding the inner tube coaxially therewith. A fuel is injected into the head portion of the inner tube and is burnt in the combustion chamber to produce hot combustion gases which are guided through a tail tube to stationary blades of a gas turbine. Combustion air is supplied through an annular passage between the inner and outer tubes from the area around the tail tube to the head portion of the inner tube. A flow-uniformalizing tube is disposed between the inner and outer tubes to cooperate with the inner tube to define therebetween an air passage.

Abstract: A combustor for a gas turbine engine including a two-stage burner having first and second combustion and exhaust sections and a burner casing coaxially surrounding the burner to define an annular conduit for reverse flow of inlet air. The burner includes a fuel injector at the upstream end thereof, primary inlet ports introducing 18% of inlet air into the first combustion section, first cooling ports introducing 12% of inlet air into the first combustion section for generating a swirling cooling flow which mixes with primary air after cooling the upstream end of the first combustion section, secondary inlet ports introducing 18% of inlet air into second stage combustion section, second cooling ports introducing 8% of inlet air into the second combustion section to generate a swirling flow which mixes with primary air after cooling the upstream end of the second combustion section, and dilution ports introducing 44% of inlet air into the exhaust section to cool the exhaust gas.

Abstract: An air storage gas turbine having a turbine, a combustion chamber supported on the external casing of the turbine and a hot gas casing leading from the combustion chamber to the blading. The hot gas casing is surrounded by an intermediate space. An inlet casing emerging into the intermediate space is provided with main air supply lines, in each of which is located at least one control device and one rapid shut-down device. The inlet casing forms an annular chamber and is provided with radial supply openings, a vertical outlet opening and axial outlet openings. A blow-out passage communicates the vertical outlet opening with a blow-out valve.